Propulsion system blade with internal actuator

ABSTRACT

Apparatus, systems, and methods are contemplated for electric powered vertical takeoff and landing (eVTOL) aircraft. Such are craft are engineered to carry safely carry at least 500 pounds (approx. 227 kg) using a few (e.g.,  2 - 4 ) rotors, generally variable speed rigid (non-articulated) rotors. It is contemplated that one or more rotors generate a significant amount of lift (e.g., 70%) during rotorborne flight (e.g., vertical takeoff, hover, etc), and tilt to provide forward propulsion during wingborne flight. The rotors preferably employ individual blade control, and are battery powered. The vehicle preferably flies in an autopilot or pilotless mode and has a relatively small (e.g., less than 45′ diameter) footprint.

This application is a continuation of U.S. application Ser. No.16/872,017 (filed May 11, 2020), which is a divisional of U.S.application Ser. No. 16/155,669 (filed Oct. 9, 2018), which is adivisional of U.S. application Ser. No. 15/985,507 (filed May 21, 2018),which claims priority to U.S. provisional application Ser. No.62/509,666 (filed May 22, 2017), U.S. provisional application Ser. No.62/509,674 (filed May 22, 2017), and U.S. provisional application Ser.No. 62/656,971 (Apr. 12, 2018), the disclosures of which are eachincorporated herein by reference.

FIELD OF THE INVENTION

The field of the invention is vertical take-off and landing aircraft.

BACKGROUND

The background description includes information that may be useful inunderstanding the present invention. It is not an admission that any ofthe information provided herein is prior art or relevant to thepresently claimed invention, or that any publication specifically orimplicitly referenced is prior art.

There is a considerable demand for electric powered transportation. Dueto the current limitations of rechargeable batteries (energy density—220Wh/Kg-300 Wh/Kg, depth of discharge, charge/discharge rates, andcycle-life issues), the general sequence of battery-powered vehicles'market entry (discounting slow short-range vehicles such as golf carts)is as follows:

-   -   Automobiles—these are the easiest vehicles to adopt electric        power. Automobiles can accept heavy batteries, there is a        relatively low battery drain rate, and operation can be safely        stopped with depletion of the battery.    -   Powered sailplanes—Here a powerplant is used for launching an        otherwise safe glider.    -   Fixed-wing training aircraft—useful for flights of short        duration, operated from established airports with professional        instructors, maintenance, and management.    -   Privately owned fixed-wing—next easiest due rolling take-off and        landing with a high wingborne lift to drag ratio.    -   Electric VTOL (eVTOL)—more challenging due the high power        required for hover, especially if high-speed efficient cruise is        also required

eVTOL has become even more challenging as the market shift from“specialized transport” aircraft making shorter trips (25-60 miles)between well-equipped terminals, to “urban mobility” aircraft makinglonger trips with at least one poorly equipped landing spot. The need isexemplified by information published by Uber™, and reproduced herein asprior art FIGS. 1A and 1B. FIG. 1A is a conceptual image of an upcomingurban transportation market for Uber®'s proposed hybrid-electricvertical takeoff and landing (eVTOL) aircraft. FIG. 1B is a projectedschedule of development and operations for such aircraft.

Safety and efficiency are perhaps the two most critical factors fordeveloping eVTOL aircraft to satisfy this market. To achieve highsafety, much of the prior art is focusing on aircraft that use six,eight, or even more, independently operated rotors. If any single rotorfails in such aircraft, the other rotors are likely to be capable ofmaking a safe landing. Even quad rotor aircraft are not considered to beparticularly fault-tolerant, because failure of a single rotor can crashthe aircraft.

Several proposed and prototype aircraft are being designed using thismany-rotor strategy. For example, FIGS. 2A and 2B are artist'srenditions of a prior art 16-rotor Volocopter™, FIG. 3 is a photographof a prior art 8-rotor Ehang™, and FIG. 4 is a photograph of a prior art8-rotor CityAirbus™. All these designs are, however, problematic becausethe rotors do not tilt from vertical lift to forward propulsionpositions, and there are no wings. That combination is extremelyinefficient in forward flight, which limits the aircraft to relativelyshort ranges.

Some eVTOL aircraft are being developed that continue to use themany-rotor strategy, but add a wing to improve forward flightefficiency. For example, FIG. 5 is an artist's rendition of a prior art36-rotor Lilium™ eVTOL, in which the rotors tilt about the forward andaft wings. The manufacturer claims a 300 km range, and 300 km/hr speed.This aircraft is, however, still problematic because the high discloading results in low power loading (high installed power per weight),which reduces efficiency and range, and produces high noise levels.

Instead of having the rotors tilt about the wings, it is possible tohave the rotors disposed in fixed position with respect to the wings,and tilt the wings. An example of that strategy is shown in FIG. 6,which is an artist's rendition of an 8-rotor Airbus™ A3 Vahana. Thisaircraft is problematic because it trades off higher efficiency inforward flight for very high power requirements during transition fromvertical lift to forward flight. In such transition, the wings act ashuge airbrakes.

It is also possible to have the rotors tilt about one or more fixedwings. Although a photograph is not available, FIG. 7 is an image of aComputational Fluid Dynamics (CFD) flow solution for the Joby™ 6-rotoreVTOL concept. This aircraft resolves some of the problems cited above,but the use of many-rotor strategy means the rotors are relativelysmall. This necessarily means high disc loading, which results in lowpower loading (high installed power per weight) and high noise level.

The only other solution that the prior art seems to have contemplated isto separate the vertical lift rotors from the forward propulsionrotors/propellers. The idea is that use of different lift and cruisepropulsion systems allows each system to be is optimized for itsparticular function. FIG. 8 is an artist's rendition of the Aurora™eVTOL concept, which uses eight lifting rotors and an aft facingpropeller. This design is problematic because the duplicate propulsionsystems require heavier and more expensive hardware, have marginal climbrate in wing borne cruise due to sizing the cruise powerplant for levelcruise, and potentially have a smaller wingborne stall speed margin—gustentry and recovery, due to design optimization for higher cruise liftcoefficient (smaller wing). Example: if flying at 130 mph, a verticalgust of 20 Ft/sec will increase the angle of attack by 6 degrees, maystall a small wing at its efficient lift coefficient of 0.9, but notstall a bigger wing at CL=0.5.

FIGS. 9A and 9B show artists' conceptions of a similar design, theTerrafugia™ eVTOL. This design has the drawbacks mentioned above withrespect to duplicate propulsion systems, and in addition, the twin tiltrotor configuration does not provide a method for pitch control in rotorborne flight.

Motor installations in the prior art are also directed towards smallrotors having small torque requirements. For example, FIG. 10 shows themotor installation of the Airbus™ A3 Vahana™. The motor is arranged in adirect-drive configuration where motor and propeller spin at the samerotational speed. This simple propulsion system solution is problematicfor large rotors with large torque requirements.

Because of the physics involved, it is relatively straightforward todesign a many-rotor eVTOL that carry a small payload (less than 500pounds) over short distances. For larger payload weights andcommercially desirable ranges the strategy of using many rotors becomesincreasingly problematic. Using a larger number of smaller rotorsprovides less disc area than fewer big rotors, requires more power peraircraft weight, is difficult to hover at low noise because the lowertotal rotor disc area results in higher blade tip Mach numbers, or inlarger number of wide-chord blades, or both, and makes autorotationflight after loss of power more dangerous, because the autorotationdescent rate increases in proportion to the square root of rotor discloading and recovery from high descent rate is risky.

Using a small number of bigger rotors (two, three or four) could solvesome of the problems discussed above, but that approach is completelycontrary to the prevailing wisdom. Among other things, thecharacteristics needed to optimize vertical lift are very different fromthe characteristics needed to optimize forward flight. In addition,those of ordinary skill in the art would dismiss the idea of havingfewer rotors on the grounds that doing so would unacceptably sacrificesafety in the event of motor failure of any of the rotors, and wouldintroduce unacceptable inefficiencies for an eVTOL. Still further, theseproblems cannot adequately be resolved by having separate lift andforward propulsion systems.

What is still needed is a vertical takeoff and landing rotorcraft thatcan safely carry a payload of at least 500 pounds, using a unitary liftand forward propulsion system with no more than four rotors, all whilepowered by the current technology battery.

All publications herein are incorporated by reference to the same extentas if each individual publication or patent application werespecifically and individually indicated to be incorporated by reference.Where a definition or use of a term in an incorporated reference isinconsistent or contrary to the definition of that term provided herein,the definition of that term provided herein applies and the definitionof that term in the reference does not apply.

In some embodiments, the numbers expressing quantities of ingredients,properties such as concentration, reaction conditions, and so forth,used to describe and claim certain embodiments of the invention are tobe understood as being modified in some instances by the term “about.”Accordingly, in some embodiments, the numerical parameters set forth inthe written description and attached claims are approximations that canvary depending upon the desired properties sought to be obtained by aparticular embodiment. In some embodiments, the numerical parametersshould be construed in light of the number of reported significantdigits and by applying ordinary rounding techniques. Notwithstandingthat the numerical ranges and parameters setting forth the broad scopeof some embodiments of the invention are approximations, the numericalvalues set forth in the specific examples are reported as precisely aspracticable. The numerical values presented in some embodiments of theinvention may contain certain errors necessarily resulting from thestandard deviation found in their respective testing measurements.

As used in the description herein and throughout the claims that follow,the meaning of “a,” “an,” and “the” includes plural reference unless thecontext clearly dictates otherwise. Also, as used in the descriptionherein, the meaning of “in” includes “in” and “on” unless the contextclearly dictates otherwise.

The recitation of ranges of values herein is merely intended to serve asa shorthand method of referring individually to each separate valuefalling within the range. Unless otherwise indicated herein, eachindividual value is incorporated into the specification as if it wereindividually recited herein. All methods described herein can beperformed in any suitable order unless otherwise indicated herein orotherwise clearly contradicted by context. The use of any and allexamples, or exemplary language (e.g. “such as”) provided with respectto certain embodiments herein is intended merely to better illuminatethe invention and does not pose a limitation on the scope of theinvention otherwise claimed. No language in the specification should beconstrued as indicating any non-claimed element essential to thepractice of the invention. Unless a contrary meaning is explicitlystated, all ranges are inclusive of their endpoints, and open-endedranges are to be interpreted as bounded on the open end by commerciallyfeasible embodiments.

Groupings of alternative elements or embodiments of the inventiondisclosed herein are not to be construed as limitations. Each groupmember can be referred to and claimed individually or in any combinationwith other members of the group or other elements found herein. One ormore members of a group can be included in, or deleted from, a group forreasons of convenience and/or patentability. When any such inclusion ordeletion occurs, the specification is herein deemed to contain the groupas modified thus fulfilling the written description of all Markushgroups used in the appended claims.

As used herein, and unless the context dictates otherwise, the term“coupled to” is intended to include both direct coupling (in which twoelements that are coupled to each other contact each other) and indirectcoupling (in which at least one additional element is located betweenthe two elements). Therefore, the terms “coupled to” and “coupled with”are used synonymously.

The following discussion provides many example embodiments of theinventive subject matter. Although each embodiment represents a singlecombination of inventive elements, the inventive subject matter isconsidered to include all possible combinations of the disclosedelements. Thus if one embodiment comprises elements A, B, and C, and asecond embodiment comprises elements B and D, then the inventive subjectmatter is also considered to include other remaining combinations of A,B, C, or D, even if not explicitly disclosed.

SUMMARY OF THE INVENTIVE SUBJECT MATTER

The inventive subject matter provides apparatus, systems and methods inwhich an electric vertical takeoff and landing (eVTOL) aircraft isengineered to carry at least 500 pounds (approx. 227 kg) using a reducednumber (e.g., 2-4) of variable speed rigid (non-articulated) rotors.

Various objects, features, aspects and advantages of the inventivesubject matter will become more apparent from the following detaileddescription of preferred embodiments, along with the accompanyingdrawing figures in which like numerals represent like components.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A is a prior art conceptual image of an upcoming urbantransportation market for Uber®'s proposed hybrid-electric verticaltakeoff and landing (eVTOL) aircraft.

FIG. 1B is a prior art projected schedule of development and operationsfor an aircraft according to FIG. 1A.

FIGS. 2A and 2B are artist's renditions of a prior art 16-rotorVolocopter™.

FIG. 3 is an artist's rendition of a prior art 8-rotor Ehang™.

FIG. 4 is an artist's rendition of a prior art 8-rotor CityAirbus™.

FIG. 5 is an artist's rendition of a prior art 36-rotor Lilium™ eVTOL,in which the rotors tilt about the forward and aft wings.

FIG. 6, which is an artist's rendition of a prior art, 8-rotor Airbus™A3 Vahana™.

FIG. 7 is an image of a Computational Fluid Dynamics (CFD) flow solutionfor the prior art Joby™ 6-rotor eVTOL concept aircraft.

FIG. 8 is a prior art artist's rendition of the Aurora™ eVTOL conceptaircraft, which uses eight lifting rotors and an aft facing propeller.

FIGS. 9A and 9B show prior art artists' conceptions of a Terrafugia™eVTOL aircraft.

FIG. 10 shows the motor installation of the prior art Airbus™ A3Vahana™.

FIG. 11 is a schematic perspective view of a preferred VTOL aircraftaccording to the inventive concepts herein.

FIG. 12 is a table of dimensions and parameters of the aircraft of FIG.11.

FIGS. 13A and 13B are schematic top and side views, respectively,dimensioned drawings of the aircraft of FIG. 11.

FIG. 13C is a table describing possible seating arrangements andcalculated weights of the aircraft of FIG. 11.

FIGS. 13D and 13E are schematic perspective views of the aircraft ofFIG. 11, with doors and hatches open.

FIG. 13F is a schematic side view of an open aft ramp of the aircraft ofFIG. 11.

FIG. 13G is a schematic view of a side section of the aircraft of FIG.11.

FIG. 14 is a schematic side view of a slotted flap that can be used withthe aircraft of FIG. 11, depicted in four different positions.

FIG. 15A is a schematic perspective view of an outboard wing foldfeature that can be used with the aircraft of FIG. 11.

FIG. 15B is a schematic front view of the outboard wing of FIG. 15A, ina folded orientation in the aircraft of FIG. 11.

FIG. 15C is a schematic top view of the aircraft of FIG. 11, with theoutboard wing in a folded orientation, such that the aircraft fits in a45′ diameter projected circular parking space.

FIG. 16A is a table describing calculated rotor geometry and propertiesas functions of non-dimensional radial station of the aircraft of FIG.11.

FIGS. 16B-16E are graphs depicting calculated beam flapwise bendingstiffness, chordwise bending stiffness, torsional stiffness and mass perunit length of the proposed rotor blades of the aircraft of FIG. 11.FIG. 16F is a graph depicting calculated center of gravity (cg) of pitchaxis versus radial station.

FIG. 17 is a graph depicting calculated lowest blade natural frequenciesat a collective control setting of zero degrees as a function of rotorspeed, of the aircraft of FIG. 11.

FIG. 18 are schematics of five cross-sectional airfoil profiles of therotor blades at specified radial stations, of the aircraft of FIG. 11.

FIG. 19A is a first schematic perspective view of the drive system thatcan be used with the aircraft of FIG. 11, enclosed in a streamlinednacelle, in position required for wing-borne flight.

FIG. 19B is a second schematic perspective view of a portion of thedrive system that can be used with the aircraft of FIG. 11.

FIG. 20 is a graph depicting calculated effects of motor speed on motorweight, for the aircraft of FIG. 11.

FIG. 21A is a schematic of a vertical cross-section of a preferredindividual blade control (IBC) configuration, which can be used with theaircraft of FIG. 11.

FIG. 21b is a schematic of a vertical cross-section of an alternativepreferred individual blade control (IBC) configuration, which can beused with the aircraft of FIG. 11.

FIG. 22 is a schematic perspective view of the aircraft of FIG. 11,having individual blade control (IBC) actuators according to FIG. 21A or21B, having four-blade primary and four-blade secondary rotors.

FIG. 23A is a schematic cross sectional view of a nacelle that can beused with the aircraft of FIG. 11, in which a battery is disposed belowthe wing, and internal to the nacelle.

FIG. 23B is a schematic cross sectional view of a nacelle and wing thatcan be used with the aircraft of FIG. 11, in which a battery is disposedwithin the wing.

FIGS. 24A-24G are schematic perspective views of an alternativepreferred VTOL aircraft according to the inventive concepts herein. Thisversion has no secondary rotors.

FIGS. 25A and 25B are schematic vertical cross-sectional views of thealternative VTOL aircraft of FIGS. 24A-24G, depicting a 3-row seatingcomparable to that of the 4-rotor configuration of FIG. 13G

FIG. 26 is a table of dimensions and parameters of the 2-rotoralternative VTOL aircraft of FIGS. 24A-24G.

DETAILED DESCRIPTION

The inventive subject matter provides apparatus, systems and methods inwhich an electric powered vertical takeoff and landing (eVTOL) aircraftis engineered to carry at least 500 pounds (approx. 227 kg) using areduced number (2-4) of variable speed rigid (non-articulated) rotors,generally assembled as primary and secondary rotors. The rotors, whetherprimary or secondary, are preferably tilt rotors such that one or moreof the rotors provides a significant amount of lift (e.g., 70%, etc)during rotor borne flight (e.g., vertical takeoff, etc), and can betilted to provide forward thrust (or air braking) during wingborneflight.

In some contemplated embodiments, each rotor can be powered by its ownelectric motor or motors, and in other contemplated embodiments,multiple rotors can be powered by a single electric motor. In especiallypreferred embodiments, individual rotors can be powered by threeelectric motors. It is also contemplated that different electric motorscould be powered by different battery packs, or multiple electric motorscould be powered by a single battery pack.

The terms “battery” and “battery pack” are used interchangeably hereinto refer to one or multiple chemical cells that produce electricity.Batteries preferably utilize Li-ion chemistries, and have a specificenergy density of about 100 kWh/lb. Other contemplated batterychemistries include Li-Polymer and Li-Metal.

Non-articulated rotors are preferred because alteration of individualblade angles can be used to apply force moments to control pitch of theaircraft in both VTOL and wingborne cruise flight. Blade angle controlis preferably achieved by individual blade control actuators preferablyfit inside their respective blades, fitted axially to the pitch axis.The individual blade control system utilized on at least each of thefirst and second primary rotors imparts a differential collective pitchbetween blades on the rotor, such that rotor thrust is maintainedapproximately constant, while shaft torque is increased above the torquerequired without differential collective. Details can be found inpending provisional applications, 62/513,930 (Tigner) “A Propeller OrRotor In Axial Flight For The Purpose Of Aerodynamic Braking”, and62/513,925 (Tigner) “Use Of Individual Blade Control To EnhanceRotorcraft Power Response Quickness”, each of which is incorporated byin its entirety reference herein.

In preferred embodiments, the aircraft has primary and secondary rotors.The primary rotors comprise blades and hubs configured to provide forforce moments at least equal to the rotor maximum lift times 6% of rotorradius, more preferably at least 9% of rotor radius, and most preferablyat least 12% of rotor radius

To achieve commercially viable flight duration, lift and othercharacteristics with no more than four rotors, and presently availablebattery technologies, at least the primary rotors needs to be relativelylarge. Accordingly, each of the primary rotors is configured to providea disc loading lower than 10 psf, and hover power loading higher than 8lb/HP. More preferably, each of the primary rotors is configured toprovide a disc loading lower than 6 psf, and hover power loading higherthan 10 lb/HP. Other contemplated aircraft embodiments have less than 8lb/HP power loading.

Furthermore, to achieve high rotor efficiency in rotor borne and inwingborne flight, a sustained rotor operation over a wide range of rotorRPM (such as 20% to 100%) is necessary, contemplated embodiments utilizerotor designs disclosed in U.S. Pat. No. 6,007,298 (Karem) “OptimumSpeed Rotor” (OSR) and U.S. Pat. No. 6,641,365 (Karem) “Optimum SpeedTilt Rotor” (OSTR).

Using the OSR and OSTR teachings, aircraft contemplated hereinpreferably achieve flap stiffness of each blade that is not less thanthe product of 100, or even more preferably 200, times the rotordiameter in feet to the fourth power, as measured in lbs-in2, at 30% ofthe rotor radius as measured from a center of rotor rotation.

Also, using the OSR and OSTR teachings, each blade weight in lbspreferably does not exceed the product of 0.004 times the diameter ofthe rotor in feet cubed.

Embodiments having first and second primary rotors are contemplated toinclude at least one optional first auxiliary rotor, each of which hasno greater than 50% of the disc area of each of the primary rotors. Inmore preferred embodiments, each of the auxiliary rotors has no greaterthan 40% of the disc area of each of the primary rotors. Auxiliaryrotors need not be the same size as each other.

The auxiliary rotor or rotors is/are also preferably rigid(no-articulated) rotors, which are configured to produce pitch forcemoments by altering the pitch of individual blades. At least the firstauxiliary rotor is advantageously configured to provide a maximumaircraft pitch force moment that is no greater than the collective totalaircraft pitch force moment capability of the primary rotors.

For each rotor, the rotating hub, the corresponding hub bearing,gearbox, and motor mounting fixture are all configured together as anintegrated rotor drive system. The preferred embodiment includes threeindependently controlled motors connected to a single gearbox perprimary rotor. The three independently controlled motors provide asafety benefit through redundancy, and additionally that configurationhas been found to be a lightweight solution for the high torque outputrequired by a variable speed rotor.

Preferred embodiments include a wing that carries at least first andsecond of the rotors, each of which is disposed in a rotor assemblyconfigured to tilt at least 90° relative to the wing. In especiallypreferred embodiments corresponding motors or other powerplants areconfigured to tilt along with rotor assemblies. At least the primaryrotors are open, i.e., as the rotors tilt they are not boundedcircumferentially by an air-directing band.

As with the rotors, the wing is relatively large relative to the weightof the aircraft and the payload. It is preferred, for example, that thewing is sized and dimensioned to provide for wing loading no higher than40 psf, and for wingborne stall speed no higher than 90 KIAS. Especiallypreferred wings are further configured to provide for wing loading nohigher than 20 psf, and for wingborne stall speed no higher than 50KIAS. Preferred wings are further configured to provide a flight speedmargin of no less than 20 KIAS in transition from fully rotor bornelevel flight to fully wing borne level flight, and wingborne cruiselift/drag ratio of no less than 10. Especially preferred wings arefurther configured to provide a transition flight speed margin of noless than 40 KIAS.

To further reduce the aircraft stall speed, the preferred wing is fittedwith an actuated slotted flap. In the especially preferredconfiguration, the flap can be used to provide aircraft roll control.The wing is preferably configured with wing tip sections having acontrol system, and electric or other actuators that adjust the wingtips to an an hedral angle of between 20-90 degrees to: (a) provide forreduced wing down load in hover; (b) provide for roll support in taxi atcross wind; and (c) provide for aircraft tie down.

The wing, rotors and other components and features discussed herein arepreferably engineered such that the aircraft can maneuver at 3 g atmaximum weight without loss of altitude or speed, but still providingthe aircraft with a low sustained autorotation descent rate if a motorfails. The preferred embodiment has a sustained autorotation descentrate of less than 1,000 ft/min.

In some embodiments at least a first battery or other power source isdisposed in the wing. Also, in some embodiments, a landing gear extendsfrom at least one of the fuselage and the wing.

In another preferred embodiment at least a first battery or other powersource is disposed in a primary rotor nacelle.

Embodiments are contemplated that have a tail and/or a canard, each ofwhich preferably has a lifting surface having an area between 10%-100%that of the wing.

Contemplated embodiments include both manned and unmanned aircraft.Thus, where a fuselage is present, it can have a passenger compartmentwith at least one seat configured to seat a human.

Electronic controls are also contemplated, sufficient to fly theaircraft without an onboard human pilot.

FIG. 11 is a perspective view of a preferred VTOL aircraft according tothe inventive concepts herein. The aircraft has a wing 1101, staticnacelle 1102, tilting nacelle 1103, fuselage 1150, tail surface 1130,and first tilting rotor system 1110. An especially preferred embodimentincludes a first tilting auxiliary rotor system 1140.

Rotor system 1110 includes rotor blades 1120. The rotor blades are of astiff hingeless variety, including for example that described in U.S.Pat. No. 6,641,365 (Karem). The rotor system collectively providesthrust as indicated by arrow 1113 and force moment 1114. The moments andforces can be controlled by rotating the blades about a feather axis1121 running the length of the blade 1120. The pitch angle around thefeather axis 1121 is represented by arrow 1122. The tip of the rotorblade follows a rotational trajectory represented by a circle 1116. Therotor blades 1120 and tilting nacelle 1103 can tilt along the pathrepresented by arrow 1111 about the tilt axis 1112. To illustrate thetilting rotor function, the right-hand nacelle is in wingborne flightorientation while the left-hand nacelle is in rotor borne flightorientation. The nacelles would be in similar orientations duringtypical operation.

Wing 1101 transmits loads from the rotor system to the fuselage 1150.Fuselage 1150 is designed to carry payload and passengers and containvarious systems including a landing gear.

FIG. 12 is a table of dimensions and parameters of a preferredembodiment, with “*” denoting turbulent flow. The preferred embodimentdescribed by the table is designed for a nominal payload ofapproximately 1,100 lbs and Basic Mission Takeoff Weight of 4,767 lbs. Awing area of 250 sq. ft. produces a wing loading of 19.1 psf. A totaldisc area of 849 sq. ft. gives a hover disc loading of 6.62 lb/ft2 whenincluding the effects of rotor wash on the airframe.

FIGS. 13A and 13B show top and side views, respectively, dimensioneddrawings of the preferred embodiment consistent with the aircraftrepresented in FIG. 11. 1310 is a main landing gear wheel attached tothe fuselage 1150. 1311 is a nose landing gear wheel attached to thefuselage 1150. All elements numbered previously are as described above.

FIG. 13B is a side view section showing the 2 upward opening nosehatches 1312, car-like 4 doors, 1313 and 1314, and baggage compartment1315 of the preferred embodiment aircraft with 3 possibleconfigurations: a) Air Taxi with one pilot and 4 passengers, b) FamilyUse with up to 8 passengers, and c) Cargo/Ambulance Use with folding aftrow seats and optional ramp replacing the baggage compartment. TheFamily Use configuration, with 1,350 Lb payload capacity, is aimed ataccommodating a family in a similar manner to that of a big SUV, exceptthat the aircraft with the 2 nose hatches provides the equivalent of 6doors as compared to the 4 doors of the SUV.

FIG. 13C is a table describing the preferred aircraft's seatingarrangement and calculated weights.

The 1,350 Lb payload capacity, the loading flexibility desired inloading the 3 rows of seating and especially the 400 Lb aft loading ofbaggage compartment or aft ramp, cause a wide aircraft C.G. shift(loading vector) of up to 8.5 inches (13.2% of wing mean aerodynamicchord). Providing aircraft stability and control with such a wide C.G.shift is made possible in rotor borne flight by the powerful pitchcontrol combination of the preferred embodiment auxiliary rotors and thepitch moment of the rigid primary rotors, and in wingborne flight by thepowerful pitch control combination of the large tail elevators and thepitch moment of the rigid primary rotors.

FIGS. 13D and 13E show views of the preferred embodiment aircraft withdoors and hatches open.

FIG. 13F shows the aft ramp 1316 open. Nose gear 1311 optionally hasvariable height, allowing adjustment of the fuselage to ground angle,providing additional clearance at the ramp opening.

FIG. 13G shows a side view section of the fuselage with a proposedseating arrangement consistent with FIG. 13C. Electronics 1317 capableof flying the aircraft without an onboard human pilot are envisioned forfuture operations.

Aerodynamic Design

Aircraft contemplated herein are designed for efficient vertical andcruise flight. Additionally, such aircraft are designed provide a safeflight and to be well-behaved in the intermediate flight conditionbetween fully wingborne and rotorborne flight known as “transition.”

Rotor thrust required for vertical flight is on the order of 10× thatrequired for efficient cruise flight. The preferred embodiment aircraftuses the variable speed rotor described in U.S. Pat. No. 6,641,365(Karem) to achieve high efficiency from 100 RPM in low speed wingborneflight to 460 RPM in hover at 12,000 feet. The rotor aerodynamic designrepresents the relatively minor compromise of optimal characteristicsfor hover vs cruise flight typical with the 5:1 RPM range available withsuch rotor. A combination of airfoil designs which have linear liftcharacteristics across a wide range of angle of attack and twist andchord distributions which balance vertical and cruise flight conditionsare required. Sectional airfoil design and analysis tools such as XFOILcan be used to design and investigate airfoils which achieve the desiredcharacteristics. Rotor analysis software for cruise such as XROTOR andsoftware for hover rotor performance such as CHARM (CDI) can be used tooptimize the rotor geometry for desired performance characteristics. Theresulting preferred rotor geometry is given as a table in FIG. 16A andairfoil sections in FIG. 18.

For efficient cruise flight, a high lift-to-drag ratio of at least 10 isdesired. The drag of the fuselage and the nacelle are minimized by usingcomputational fluid dynamics (CFD) programs, for example STAR-CCM+, toanalyze and iteratively optimize the shape subject to practicalconsiderations such as volume for propulsion and payload and forstructural requirements. The wing airfoil can be optimized using airfoiltools such as the aforementioned XFOIL. Considerations pertinent to wingoptimization include a compromise between cruise drag, download invertical flight, maximum lift in transition, and structuralrequirements.

A preferred method for increasing maximum lift in transition withoutnegatively affecting cruise drag is a slotted flap, as shown in thesectional drawing FIG. 14. The figure includes multiple flap deflectionangle positions: maximum up deflection (−8), no deflection (0), maximumCL (+22), and maximum down deflection (+65). The flap 1401 rotates abouta simple hinge 1404. When are tracted (position 0), the flap 1401 incursminimal additional drag compared to a single element airfoil. Whendeployed to an optimal high lift angle (position+22), a slot is exposedwhich allows airflow from the lower side of the first element 1402 topass over the flap element 1401. In the preferred embodiment, the flapand slot shape are designed to provide linear lift response to smalldeflection angles such that the slotted flaps are additionally capableof precise control as ailerons for aircraft roll control. The preferredslotted flap extends spanwise from the fuselage side to the wing tipwith an interruption at the propulsion system nacelle. The flap issegmented into multiple spanwise sections to reduce stresses induced bywing deformation. The flap hinge is offset from the wing surface suchthat a slot opens as the flap is deflected downward. A flexible uppersurface seal 1403 minimizes drag when the flap is in its retractedposition.

The most flight safety critical phase of a winged eVTOL flight is thetransition from fully rotor borne to wing borne at a safe forward speed.This is especially important in turbulent windy conditions, in lowaltitude urban environments. Unlike the prior art, the use of largeprimary rotors and large wing area combine to provide safe transition.

In the preferred embodiment (with 2 auxiliary rotors), the use of largerotors results in disc loading, which produces low noise (350 RPM, rotortip Mach number lower than 0.35), efficient steady hover, and 2 g rotorborne maneuvering at 495 RPM. In combination with a large (250Ft{circumflex over ( )}2) wing having a slotted flap at 22 degrees, theaircraft is engineered to provide a stall speed of 50 KIAS at anaircraft weight of 4,767 Lb. At rotor speed of 550 RPM the aircraft canbe fully rotor borne (zero wing and tail lift) at 90 KIAS, 40 KIAShigher than the minimum wingborne speed, and can carry 2.5 ginstantaneous rotor lift. At 90 KIAS the aircraft can have 3.25 gwingborne lift. These large margins avoid most flight accidents typicalof low lift and control accidents typical of low speed flight andtransition in turbulent weather

Folding Outboard Wing

The outboard wing fold feature is depicted in FIG. 15A. The outboardwing 1501 folds with respect to inboard wing 1502 about hingeline 1503,and the motion is controlled by a fold actuator (not shown). The foldingwing actuation is designed to withstand flight and ground loads. Itfeatures a sprung skid 1504 at the wingtip which contacts the groundupon landing. The large wingspan necessary for efficient flight andacceptable transition characteristics makes the aircraft sensitive tocross wind and wind gust on the ground. Tip skids give the aircraftadditional ground stability and safety. The wingtip also includestie-down features (not shown) for securing the aircraft while parked. Inrotor-borne flight, the download due to rotor wash on the outboard wingis reduced because download is inversely related to rotor-wingseparation distance. FIG. 15B shows the front view of the aircraft withthe outboard wings folded. Additionally, the folded aircraft can fit ina smaller parking area. FIG. 15C shows the top view of the aircraft,which fits in a 45′ diameter projected circular parking space.

Blade Design

FIG. 16A gives the rotor geometry and properties as functions ofnon-dimensional radial station. FIG. 18 shows cross-sectional airfoilprofiles of the rotor blade at specified radial stations.

FIG. 16B depicts the flapwise, or normal to the chord, bending stiffnessof the exemplary embodiment rotor blade from radial station zero at theroot to radial station 1 at the tip. FIG. 16C depicts the lag, orchordwise, bending stiffness of the exemplary embodiment rotor bladefrom radial station zero at the root to radial station 1 at the tip.FIG. 16D depicts the torsional stiffness of the exemplary embodimentrotor blade from radial station zero at the root to radial station 1 atthe tip. FIG. 16E depicts the mass per unit length of the exemplaryembodiment rotor blade from radial station zero at the root to radialstation 1 at the tip. FIG. 16F depicts the chordwise cg positionrelative to the blade pitch axis of the exemplary embodiment rotor bladefrom radial station zero at the root to radial station 1 at the tip.

FIGS. 16C-E depict the beam flapwise bending stiffness, chordwisebending stiffness and torsional stiffness of the proposed embodimentblades. High stiffness to mass is required to avoid structural dynamicsproblems while operating the rotor over a large range of rotor speeds.The blade mass distribution of the proposed embodiment blade is shown inFIG. 20F. To avoid aeroelastic instabilities, the chordwise center ofmass of the rotor blade cannot be much further aft than the blade pitch,or feather, axis. FIG. 16F is a graph depicting calculated center ofgravity (cg) of pitch axis versus radial station. The embodiment rotorblade was found to be free of aeroelastic instabilities over theoperating conditions with the center of mass balanced as depicted inFIG. 16F.

FIG. 17 depicts the lowest blade natural frequencies at a collectivesetting of zero as a function of rotor speed. The rays emanating fromthe origin depict the 1/rev, 2/rev . . . 10/rev harmonic frequencies ofthe rotor. Operating speed ranges are marked on the horizontal axis. Thestiffness and mass distributions of FIGS. 16A-F are the primaryinfluences on the rotor blade natural frequencies. The naturalfrequencies remain well separated from each other across the operatingrange. As a result of the high stiffness to weight design, the naturalfrequencies are much higher than typical rotor blades. The first flapmode remains above the 3/rev excitation frequency of the rotor acrossthe entire operating range whereas the more lightly damped first lagmode remains above the 4/rev excitation frequency. This separation abovethe primary rotor excitation frequency of 3/rev permits operation over awide rotor speed range without encountering excessive vibratory loads orvibrations due to resonance.

Rotor dynamics simulation and optimization software programs such asCHARM and CAMRAD may be used to iterate the rotor blade design subjectto the desired characteristics described. Finite Element Analyses (FEA)software may be used for higher fidelity structural analysis, and CFDcodes may be used for higher fidelity aerodynamic analysis andrefinement.

The preferred auxiliary rotor and blades are designed following the sameperformance constraints as the primary rotor with a smaller diameter.

Hub Drive System

The drive system is enclosed in a streamlined nacelle, illustrated inFIG. 19A in the position required for wing-borne flight. The axis ofrotation of the rotor is show as X-X and the direction of flight byArrow A. Tilt rotor aircraft, by definition, require the thrust axis ofthe rotor to be rotated from the horizontal flight condition to thevertical lift condition. This angle is not less than 90 Deg. and can be105 Deg. or more. The axis about which the forward section of thenacelle tilts, including all drive elements, is shown as Y-Y.

Blade Shanks, 1901, are mounted in Feather Bearing Containment Hoops,1902, which are bolted to the rotating Hub, 1903 supported on largediameter Bearing, 1904. The three Motors, 1905, are symmetricallydisposed about the hub center, one of which is shown sectioned, 1906.The output Sun Gear, 1907, is driven via Sprag Clutch, 1908. The PlanetGears, 1909 are mounted in Planet Carrier, 1910 which is attached toOutput Pinion, 1911. The three identical output pinions mesh with RingGear, 1912. Hub loads are carried from the hub bearing throughIntermediate Structure, 1913 which is attached by bonding and rivetingto Nacelle, 1914, of monocoque composite construction. The shellstructure is attached to the aft nacelle at Hinge Points, 1915, with thetilt actuation Truss, 1916, connecting both nacelle elements at actuatorAttachment Bracket, 1917. The electronics motor Driver Boxes, 1918, areindividually packaged for redundancy, with Phase Connections, 1919 tothe motors. The motor Liquid Cooling connection, 1920, is illustrated,as is the Oil Containment Sump, 1921. The alternative Rotary TiltActuator, 1922, is shown mounted on a transverse axis of rotation.

FIG. 19B depicts an alternative perspective view of a portion of thestreamlined nacelle depicted in FIG. 19A. FIG. 19B provides a closerview of Blade Shanks 1901 as mounted in Feather Bearing ContainmentHoops 1902, as well as the bolting of Feather Bearing Containment Hoops1902 to rotating Hub 1903, which is supported on large diameter Bearing1904.

The entire rotor hub, including the blade feather bearings and pitchactuation system coupled with the electric drive, form an integratedassembly. The system is illustrated as a three-bladed arrangement; otherblade numbers are similarly installed. The four predominating loadsresolved through the assembly from the rotating frame to the nacellestructure are the blade flap loads, the mast moment, the thrust or liftvector and the drive torque. A large diameter, moment-carrying bearingconnects the rotating hub elements to the nacelle structure. The large,slow-turning rotor creates a drive condition where the rotortorque/speed characteristics are well beyond the capabilities of adirect-drive motor. An analysis summary of the effects of motor speed onmotor weight is shown in FIG. 20, indicating that for constant power,the high-speed motor with gear reduction offers weight reduction withincreasing gear ratio, and, further weight reduction is realized bymultiplexing the motors. Because flight safety is of prime importance,multiplexed motors offer complete electrical redundancy. A degree ofmechanical redundancy is also provided by the inclusion of a one-wayclutch (“sprag” clutch) on each of the motor output shafts. The systemillustrated has three motors but is adaptable to a larger complement ofmotors.

The typical application of a direct drive electric motor for eVTOL liftrotors has several advantages. It is simple, and with simplicity comesinherent reliability. Also, it avoids the additional weight of agearbox. However, a weight reduction is also available by increasing theelectric motor RPM and gearing the output stage. By trading torque forRPM at fixed power, significant weight savings are available. There is alimit to the weight reduction based on the limited ability for coolingas size decreases. Therefore, for weight optimization, the choicebetween a direct drive and geared motor depends on the desired poweroutput and desired RPM. At lower output RPM a geared drive provides aweight advantage and at higher RPM a direct drive is more weightefficient. FIG. 20 shows the weight of the motors and drive of the 300HP rated output of the primary rotor in the preferred embodiment. Inthis case the weight breakpoint is near 2000 RPM. In the preferredembodiment the primary rotor RPM in high-power hover is 400-460 and theRPM in high-power wingborne climb is 350, resulting in a clearpreference of a geared rotor drive.

There are several advantages to incorporating more than one motor in thedesign. A) More motors can be more weight efficient. The ability todissipate heat scales with the surface area of the motor and the power,for fixed RPM, scales with the volume, and therefore the weight, of themotor. A greater surface area to volume ratio allows for better cooling.In the case that the minimum weight is determined primarily by theability to cool, more motors provide better weight efficiency becausethey have a higher surface area to volume ratio. B) Reliability can beincreased through redundancy. In a configuration in which acceptableoutput power can be maintained with one or more drive motors failed,overall reliability is increased through redundancy. However, thecomplexity of many motors can reduce reliability.

The higher the gear ratio the better the weight savings. Because thegearbox weight is driven by the high-torque output stage, to firstorder, the gearbox weight is independent of the gear ratio. As shown inFIG. 20, the weight of the motors scale inversely with RPM. However, theweight savings is limited by the reduced ability for heat dissipation assize decreases. In addition, there are practical limitations to motorRPM that include: retention of magnets under high centrifugal force (formotors of that design type), limitations in available bearing speeds,and limitations of electronic switching speeds for motor commutation.

In the current example a gear ratio of 20:1 provides weight reductionwhile limiting the challenges of very high motor RPMs. In othercontemplated embodiments, the gearbox may have ratios of 3:1, 5:1, 10:1,20:1, or 30:1.

With an overall gear ratio close to or exceeding 20:1, two stages ofgear reduction are required for the full weight-saving benefits ofhigh-speed motors to be realized. Each motor is equipped with aplanetary reduction set driving a combining ring gear attached to thehub. All features of the assembly are optimized for minimum weight, forexample, the use of three driver pinions engaging with a single largering gear minimizes the face width of the ring with consequent materialsaving.

The motors, their driver electronics and the gearbox in total requirecooling. The preferred fluid for motor and electronics cooling iswater/glycol, and a further liquid-to-oil heat exchanger is employed forgearbox oil cooling. Gearbox oil is contained in a sump located at thelower aft extremity of the gearbox housing.

The nacelle tilt system is shown as system of three linear actuators,the aft. pair providing 60 Deg. of nacelle travel and the fwd. actuatorthe remaining 55 Deg. An alternative system is the application of ahigh-torque rotary actuator operating through a four-bar linkage. Therotary actuator, U.S. Pat. No. 7,871,033 (Karem et al.) which describesits execution in detail, is cited in the references.

Individual Blade Control

In the preferred embodiment, individual blade control (IBC) actuators2101 enable precise, independent control of the rotor bladetrajectories. By independently controlling the blade angle, the rotormoment and forces can be controlled. FIG. 22 shows a similar to thatdepicted in FIG. 11, except that first tilting rotor system 2210 andfirst tilting auxiliary rotor system 2240 are 4-blade rotorapplications. Application of IBC actuation with a 4-blade rotor enablesaircraft braking (negative rotor thrust) in wingborne flight withoutintroducing large hub moments. All elements similarly numbered as inFIG. 11 are described as above. Details can be found in pendingprovisional applications, 62/513,930 (Tigner) “A Propeller Or Rotor InAxial Flight For The Purpose Of Aerodynamic Braking”, and 62/513,925(Tigner) “Use Of Individual Blade Control To Enhance Rotorcraft PowerResponse Quickness”, each of which is incorporated by in its entiretyreference herein.

A preferred IBC configuration is shown in FIG. 21A. The design approachis to locate an electric actuator within the blade itself, positioned sothe actuator and blade feather axes are concurrent. Certain blade designconditions have to be met for this approach to be feasible. OSTR rotorblades should, with great advantage, have a high stiffness in flapbending and lead-lag which leads to blade root section chords andthicknesses much larger than seen in conventional rotor blades. Theresultant blade spar, being hollow and of adequate diameter,conveniently accepts the cylindrical electric actuator. In combinationwith a reduction gearbox, it will be seen that electric motor drive canconnect the blade to the hub in a rotational sense without the need formechanical links and can be commanded and controlled exactly as otherflight control actuators. The general term applied to this type ofactuation is Individual Blade Control (IBC) which permits an entirelynew and optimized matrix of blade azimuth and pitch angle. There areaerodynamic advantages in doing so.

Additionally, military helicopters engaged in ship-borne operations haveto be made compact by means of folding. Folding an existing state of theart rotor blade and maintaining the integrity of the pitch linkageresults in a complex arrangement of mechanical parts. The subjectinvention eliminates this complexity; and the only new requirement forthe actuator design, being internal to the blade, is that the electricalcabling flexes with the fold angle. This requirement is readily andsimply achieved.

Hollow Blade Spar, 2101 is inserted into receiving bore of Hub, 2102supporting the blade by Bearing, 2103 running on Inner Race, 2104 andsealed by Seal, 2105. If required to fold, the blade and hub portionrotates about Hinge, 2106. Either one Motor Stator, 2107, or, two MotorStators, 2107 and 2108 operate Rotor, 2109, guided by Tail Bearing, 2110and Rotor Bearing, 2111. The motor rotors position and hence bladeangular position is sensed by Encoder, 2112 with a static reference bymeans of Stationary Core, 2113. The motors drive the Gearbox, 2114,which is secured to the blade root by Fastenings, 2115. The gearboxreaction torque is carried by Flexible Coupling, 2116, whose purpose isto isolate the gearbox from moment-induced deflections resulting fromblade flap and lead-lag loads. Centrifugal loads as well asmoment-induced radial loads are carried by Taper Roller Bearing, 2117.Blade actuation torque is reacted through Spline, 2117 and theCentrifugal Force is reacted by Nut, 2119. Flexible electricalConnection Cable, 2120 carries motor power and control information fromSlip Ring, 2121, which rotates about Hub Rotation Axis, 2122 (shown byAxis X-X) with the static portion of the slip ring supported by AirframeStructure, 2123. The Coolant Fluid flow and return lines, 2124, are fedthrough Rotary Gland, 2125.

FIG. 21B shows an alternative component layout. If the rotor blade isnot required to fold, and the blade feather axis is tightly controlledrelative to the hub by means of rigid feather bearings, then the pitchactuator can be hub-mounted as opposed to blade-mounted. In thisarrangement, the actuator assembly consisting of motor or motors, thereduction gearbox and the required sensors and connection wiring isconnected to the hub. This connection is torsional stiff, but flexiblein alignment to accept the deflections inherent in highly-loaded blades.The splined output drive disc mates with an engaging spline internal tothe blade. There are practical system advantages in co-mounting a set ofactuators on a unitary hub, as they can be electrically connected withcommon driver, power, and cooling paths.

Cylindrical blade spar, 2131, is supported in the Outboard FeatherBearing Assembly consisting of Bearing Retainer Hoop, 2132, Outer Race,2133, Rollers and Cage, 2134, Inner Race, 2135, and Seals, 2136. Theblade root is stabilized with Inner Diaphragm, 2137, secured withRivets, 2138. The diaphragm is internally splined at 2139 for torquetransfer from the Flexible Drive Bellows, 2140. This is the point ofseparation when the removed blade is withdrawn over the fixed actuator.

The Split Blade Retaining Clamp, 2141, secures the Inboard Root Fitting,2142, to the Inboard Feather Bearing Outer Race, 2143. The Taper Rollerand Cage, 2144, runs on Inner Race, 2145, sealed by Seal, 2146. Bearingpre-load is provided by Belleville Spring, 2147, operating on ThrustWasher, 2148.

Blade centrifugal force is reacted by Actuator Housing, 2149, retainedby Fastener Set, 2150 which also secures Static Core, 2151 to theRotating Hub Component, 2152. The static core carries both Motor StatorWindings, 2153 and the Position Encoder, 2154. The electric motor Rotor,2155 is supported in Journal Bearing, 2156 and Tail Bearing, 2157, anddrives the Reduction Gearbox, 2158.

A tubular Extension, 2159, attached to airframe structure, non-rotating,carries Slip Ring, 2160, providing current and control signals to theactuator via fixed wiring Harness, 2161.

Battery

The preferred battery installation is shown in the nacelle crosssectional view in FIG. 23A. The battery 2301 is disposed below the wing2302 and internal to the nacelle 2303. Flight direction is indicated byblock arrow B. The aft volume of the nacelle includes sufficient volumeto include portions of the cooling system required for the electricpropulsion system. Alternatively, the battery 2311 can be contained inthe wing structure 2312 as shown in FIG. 23B. The battery 2311 is shownof smaller cross-sectional dimensions than in FIG. 23A but it providesthe same volume because it is enclosed by the long wing 2321. Thenacelle 2313 can then be significantly smaller and lower drag than inthe primary preferred embodiment. Alternatively, a nacelle arrangementwith an internal combustion engine and a generator (hybrid propulsion)can be used to power the electric motors and other aircraft systems andprovide for a substantially longer range.

Alternative Configuration

FIGS. 24A-G show an alternative preferred embodiment without auxiliaryrotors. This preferred embodiment uses the powerful pitch control of theprimary rotors 2401 in rotor borne flight and the pitch control of along control arm canard surface 2411 in wingborne flight to makepossible an aircraft with lower weight, drag, installed power and cost,providing the same aircraft performance but with reduced aircraftcontrol in transition and in wind gust and reduced level ofaccommodation of C.G. shift and payload versatility (no cargo ramp andreduced luggage volume and weight) as compared to the preferredembodiment with the auxiliary rotors and large tail area.

The alternative preferred embodiment features a fuselage 2421, wing2431, rotor blades 2401, and canard 2411. The internal configuration issimilar to the primary preferred embodiment. The fuselage has 3 rows ofseating: front row 2501, middle row 2502, and aft row 2503.

While the cabin volume of the 2-rotor alternative configuration iscomparable to that of the one with the 4 rotors, its wingborne drag isreduced by: a) wing area reduced from 250 Ft{circumflex over ( )}2 to140 Ft{circumflex over ( )}2, b) no tail section, c) no auxiliary rotornacelles, d) wing to fuselage attachment behind the cabin (lower frontalarea), e) option for extensive fuselage laminar flow, and f) lowercruise drag due to lower cruise weight (estimated as 817 Lb lighter dueto a lighter airframe and smaller battery).

FIGS. 25A and 25B show the inboard profile of the 2-rotor alternativeconfiguration demonstrating a 3-row seating comparable to that of the 4rotors configuration. FIG. 26 is a table of dimensions and parameters ofthe 2-rotor alternative configuration.

MODIFICATIONS

It should be apparent to those skilled in the art that many moremodifications besides those already described are possible withoutdeparting from the inventive concepts herein. The inventive subjectmatter, therefore, is not to be restricted except in the spirit of theappended claims.

What is claimed is:
 1. An aircraft rotor system comprising: a rotorblade pitch actuator; a rotor blade, wherein the rotor blade pitchactuator is internal to the rotor blade.
 2. The aircraft rotor system ofclaim 1 wherein the aircraft rotor is a rigid rotor.
 3. The aircraftrotor system of claim 1 wherein the rotor blade pitch actuator comprisesan electric actuator.
 4. The aircraft rotor system of claim 3 whereinthe electric actuator comprises an electric motor rotor and the electricmotor rotor is fixed relative to the rotor blade.
 5. The aircraft rotorsystem of claim 1 wherein the rotor blade pitch actuator comprises afirst and second motor stator internal to the rotor blade.
 6. Theaircraft rotor system of claim 1, wherein the rotor blade pitch actuatoradditionally comprises a gearbox, wherein the gearbox is connected to arotor hub and connected to the rotor blade.
 7. The aircraft rotor systemof claim 1 additionally comprising a flexible coupling that is torqueconnected to the blade and to the rotor hub.
 8. An aircraft rotorcomprising a rotor blade having a blade pitch axis, and a first and asecond rotor blade pitch actuators disposed concentric with the bladepitch axis.
 9. The aircraft rotor of claim 8 wherein the aircraft rotoris a rigid aircraft rotor.
 10. The aircraft rotor of claim 8 wherein thefirst blade pitch actuator comprises an electromagnetic actuator. 11.The aircraft rotor of claim 10 wherein the electromagnetic actuatorcomprises an electric motor rotor and the electric motor rotor is fixedrelative to the rotor blade.
 12. The aircraft rotor of claim 8additionally comprising a gearbox, wherein the gearbox is connected to arotor hub and connected to the rotor blade.
 13. The aircraft rotor ofclaim 8 additionally comprising a flexible coupling that is torqueconnected to the blade and to the rotor hub.
 14. An aircraft comprising:a propulsion system blade; an axially aligned propulsion system bladepitch actuator wherein the propulsion system blade pitch actuator isconfigured to actuate the propulsion system blade without mechanicallinkages.
 15. The aircraft of claim 14 wherein the propulsion systemblade pitch actuator comprises a gear reduction system.
 16. The aircraftof claim 14 wherein the propulsion system blade pitch actuator comprisesan electric motor.
 17. The aircraft of claim 16 wherein the electricmotor comprises an electric motor stator that remains fixed relative toa propulsion system hub.
 18. An aircraft propulsion system comprising: apropulsion system blade pitch actuator; a propulsion system blade,wherein the propulsion system blade pitch actuator is internal to thepropulsion system blade.
 19. The aircraft propulsion system of claim 18wherein the rotor blade pitch actuator comprises an electric actuator.20. The aircraft propulsion system of claim 18 wherein the rotor bladepitch actuator comprises a first and second motor internal to the rotorblade.